In the compressor portion of an aircraft gas turbine engine, atmospheric air is compressed to 10-25 times atmospheric pressure, and adiabatically heated to 800°-1250° F. in the process. This heated and compressed air is directed into a combustor, where it is mixed with fuel. The fuel is ignited, and the combustion process heats the gases to very high temperatures, in excess of 3000° F. These hot gases pass through the turbine, where rotating turbine wheels extract energy to drive the fan and compressor of the engine, and the exhaust system, where the gases supply thrust to propel the aircraft. To improve the efficiency of operation of the aircraft engine, combustion temperatures have been raised. Of course, as the combustion temperature is raised, steps must be taken to prevent degradation of engine components directly and indirectly as a result of the higher operating temperatures.
The requirements for enhanced performance continue to increase for newer engines and modifications of proven designs, as higher thrusts and better fuel economy are among the performance demands. To improve the performance of this engine, the combustion temperatures have been raised to very high temperatures. This can result in higher thrusts and/or better fuel economy. These combustion temperatures have become sufficiently high that even superalloy components not within the combustion path have been subject to degradation. These superalloy components have been subject to degradation by mechanisms not generally experienced previously, creating previously undisclosed problems that must be solved. One recent problem that has been discovered during refurbishment of high performance aircraft engines has been the pitting of turbine disks, seals and other components that are supplied with cooling air. The cooling air includes ingested particulates such as dirt, volcanic ash, fly ash, concrete dust, sand, sea salt as well as metal, sulfates, sulfites, chlorides, carbonates, various and sundry oxides and/or various salts in either particulate or gaseous form. These materials are deposited on substrate surfaces. When deposited on metallic surfaces, these materials can interact with one another and with the metallic surface to corrode the surface. Corrosion is accelerated at elevated temperatures. The materials used in turbine engines are typically selected on high temperature properties, including their ability to resist corrosion, even these materials will degrade under severe conditions at elevated temperatures. On investigation of the observed pitting problem, it has been discovered that the pitting is caused by a formation of a corrosion product as a result of the ambient airborne foreign particulate and gaseous matter that is deposited on the disks, seals or other components as a result of the flow of cooling air containing foreign particulate and gaseous matter. This deposition, along with the more elevated temperature regimes experienced by these engine components, has resulted in the formation of the corrosion products. It should be noted that the corrosion products are not the result of exposure of the engine components to the hot gases of combustion, normally associated with oxidation and corrosion products from contaminants in the fuel. The seals, turbine disks and other components under consideration and discussed herein generally are designed so that, if a leak is present, the air will leak in the direction of the flow of the hot gases of combustion and not in the direction of the components under consideration.
Because the corrosion products are the result of exposure of the engine components to cooling air drawn from ambient air environments, it is not uniform from engine to engine as aircraft visit different geographic locations with different and distinct atmospheric conditions. For example, some planes are exposed to salt water environments, while others may be subject to air pollutants from highly industrial regions. The result is that some components experience more advanced corrosion than other components.
The corrosion was not unanticipated. But the remedial efforts initiated during the production were ineffective. Various coatings have been suggested and attempted to mitigate corrosion concerns. One is a phosphate-based set forth in U.S. patent application Ser. No. 11/011695 entitled CORROSION RESISTANT COATING COMPOSITION, COATED TURBINE COMPONENT AND METHOD FOR COATING SAME filed on Dec. 15, 2004, assigned to the assignee of the present application and incorporated herein by reference. Others include aqueous corrosion resistant coating compositions comprising phosphate/chromate binder systems and aluminum/alumina particles. See, for example, U.S. Pat. No. 4,606,967 (Mosser), issued Aug. 19, 1986 (spheroidal aluminum particles); and U.S. Pat. No. 4,544,408 (Mosser et al), issued Oct. 1, 1985 (dispersible hydrated alumina particles). Corrosion resistant diffusion coatings can also be formed from aluminum or chromium, or from the respective oxides (i.e., alumina or chromia). See, for example, commonly assigned U.S. Pat. No. 5,368,888 (Rigney), issued Nov. 29, 1994 (aluminide diffusion coating); and commonly assigned U.S. Pat. No. 6,283,715 (Nagaraj et al), issued Sep. 4, 2001 (chromium diffusion coating). A number of corrosion-resistant coatings have also been specifically considered for use on turbine disk/shaft and seal elements. See, for example, U.S. Patent Application 2004/0013802 A1 (Ackerman et al), published Jan. 22, 2004 (metal-organic chemical vapor deposition of aluminum, silicon, tantalum, titanium or chromium oxide on turbine disks and seal elements to provide a protective coating). These prior corrosion resistant coatings can have a number of disadvantages, including: (1) possibly adversely affecting the fatigue life of the turbine disks/shafts and seal elements, especially when these prior coatings diffuse into the underlying metal substrate; (2) potential coefficient of thermal expansion (CTE) mismatches between the coating and the underlying metal substrate that can make the coating more prone to spalling; and (3) more complicated and expensive processes (e.g., chemical vapor deposition (CVD)) for applying the corrosion resistant coating to the metal substrate.
Still another problem is that a corrosion mitigation coating that has been applied to certain components has proven to be ineffective. This coating, an alumina pigment in a chromate-phosphate binder utilizing hexavalent chromium in a coating composition commercially marketed as SermaFlow® N3000, cracked after exposure to elevated temperatures. SermaFlow® is a registered trademark of Sermatech International of Pottstown, Pa., USA. Of course, that coating also has the disadvantage of including the environmentally unfriendly element, chromium, which presents challenges during application. Additionally, while such a coating is effective at low temperatures, it has a low coefficient of expansion so that at the higher temperatures experienced by newer engines, the coating, even when applied in thicknesses of as thin as 0.5-2.5 mils, cracked. In fact at thicknesses of 1.5 mils and greater, this coating delaminated after one engine cycle at 1300° F., a capable operating temperature for newer engines. While the problem described has been most evident on the newer high performance engines, because of the extremes dictated by its operation, the problem is not so restricted. As temperatures continue to increase for most aircraft engines as well as other gas turbine engines, the problem will also be experienced by these engines as they cross a temperature threshold related to the materials utilized in these engines.
What is needed is a coating composition that is free of hexavalent chromium that can be applied to prevent corrosion of turbine engine components even when the turbine engine components are subjected to elevated operating temperatures in a wide variety of atmospheres.